1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with a cooling circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Turbine airfoils, such as rotor blades and stator vanes, pass cooling air through complex cooling circuits within the airfoil to provide cooling from the extreme heat loads on the airfoil. A gas turbine engine passes a high temperature gas flow through the turbine to produce power. The engine efficiency can be increased by increasing the temperature of the gas flow entering the turbine. Therefore, an increase in the airfoil cooling can result in an increase in engine efficiency.
Prior art airfoil cooling of blades with serpentine airfoil cooling circuits allows for the cooling air to communicate in between the mainstream pressure side and suction side. This cooling circuit design has to compromise the mainstream heat load and pressure distribution on the airfoil pressure and suction walls. FIG. 1 shows a prior art serpentine flow cooling circuit with a cooling cavity to provide cooling air for both the pressure and suction sides of the blade. A leading edge of the blade is cooled with a showerhead arrangement in which cooling air is supplied through a leading edge cooling supply channel 11, passes through a plurality of metering holes 12 and into a leading edge cavity 13, and then the cooling air is discharged out film cooling holes 14 that form the showerhead and gill holes. A mid-chord region of the blade is cooled by cooling air supplied through a first leg 15 of a three-pass serpentine forward flow circuit, and flows through the serpentine path into the second leg 16 and the third leg 17 in a forward direction from the trailing edge to the leading edge of the blade. Blade tip exit holes 23 also discharge cooling air from the serpentine flow circuit and out through the blade tip to provide cooling thereof. The first leg 15 of the serpentine flow circuit passes the cooling air through a series of three impingement holes 18, 19, 20 formed along the trailing edge of the blade before exiting out exit cooling air holes 21 spaced along the trailing edge of the blade. Film cooling holes 22 are located along the pressure side and suction side of the blade and connected to the first leg 15 of the serpentine flow circuit to provide film cooling to the outer surface of the blade. FIG. 1 also shows a schematic diagram representing the cooling air flow paths through the blade in FIG. 1.
U.S. Pat. No. 7,033,136 B2 issued to Botrel et al on Apr. 25, 2006 entitled COOLING CIRCUITS FOR A GAS TURBINE BLADE discloses a gas turbine blade best seen in FIG. 4 of this patent. The blade includes a first admission opening and a second admission opening formed in the root of the blade to supply pressurized cooling air to the blade cooling circuit. Cooling air from the admission openings flow into the suction side cavity or the pressure side cavity along the spanwise direction of the blade. The cooling air in these side cavities then flows into a common central cavity extending radially in the central portion of the blade between the suction side cavity and the pressure side cavity. According to FIG. 2 of this patent, two rows of film cooling holes are connected to the central cavity to discharge film cooling air onto the pressure side surface of the blade. One major difference between the Botrel patent and the present invention is that the pressure and suction side cavities are cavities and not individual radial channels. As such, the channels cannot be individually sized such that specific pressure and flow can be designed depending upon the hot metal temperature occurring on the blade. Another major difference is the use of a common central cavity used for both the pressure side supply cavity and the suction side supply cavity. Both supply cavities discharge into the common central cavity. In the present invention, separate collector cavities are used, one for the pressure side supply channels and one for the suction side supply channels. The use of separate collection cavities for the pressure and suction sides allow for better control of the pressure and flow distribution of the cooling air around the sections of the blade.
The object of the present invention is to provide for a turbine blade with multiple individual zones having independent designs based on the local heat load and aerodynamic pressure loading conditions.
Another object of the present invention is to provide for a turbine blade with near wall cooling so that the airfoil can be made thin to increase the airfoil overall heat transfer convection capability.
Still another object of the present invention is to separate the pressure side flow circuits from the suction side flow circuits in order to eliminate back flow margin design issues and high blowing ratio for the airfoil suction side film cooling holes.